The present invention relates generally to thermal insulation systems and specifically to external insulation for carbon/carbon combustion chambers of ramjet-powered missiles.
Future long-range missiles are expected to employ ramjet engines operating combustion temperatures of 3000.degree. F. or higher. Carbon/carbon is a prime material candidate for these ramjet combustors because of its high-temperature operational capability. However, adjacent missile equipment and structure must have some form of thermal protection from the high combustor wall temperatures.
Additionally, some future long-range missiles are expected to employ integral-rocket ramjet engines. This engine concept consists of a solid rocket motor case which also serves as a ramjet combustion chamber. During the initial boost phase, when the solid rocket fires, internal pressures are high and wall temperatures are low. Later, during the ramjet phase, internal pressures are low but wall temperatures are high. Carbon/carbon has enough strength for the high boost pressures and can withstand temperatures as high as 5000.degree. F.
As mentioned above the hot wall combustor must be insulated externally to protect adjacent equipment and structure. A particular prior art thermal protection system (TPS) uses the carbon/carbon combustor wall as one of three layers of different insulation materials for ramjet combustion chambers. Combinations of different insulation materials are often used in separate layers to taylor the thermal protection for the different useful temperature limits of the materials, for changing thermal conditions during a flight as well as for varied flight times. In this prior art TPS, the temperature of each insulation layer is maintained below its useful temperature limit, and no thermal degradation of any material is permitted.
The first layer of the prior art insulation system is the combustor external wall formed from a carbon/carbon composite. Carbon/carbon composites provide ideal combustion chambers since they can withstand temperatures of 5,000.degree. F. and high boost pressures. However, the thermal conductivity of carbon/carbon is relatively high resulting in wall temperatures approaching the combustion gas temperature at steady-state conditions.
The second layer of the prior art insulation system is a layer of zirconia. This insulation layer is deposited upon the external wall of the carbon/carbon combustor. Zirconia has a 4000.degree. F. useful temperature limit which is less than that of carbon/carbon. However, zirconia has a lower thermal conductivity than the carbon/carbon material. The zirconia insulation reduces the hot wall combustor temperature down to the useful temperature limit of the outer insulation layer.
The outer layer of the prior art system is known as "MIN-K 2000", a proprietary insulation material of the Manville Corporation. "MIN-K" is a flexible blankoet material with a quartz cloth facing. This outer insulation layer is used as an intermediate-temperature (2000.degree. F.) insulator between missile equipment and the zirconia layer. "MIN-K" is an ideal thermal insulation because of its low thermal conductivity which is significantly below that of zirconia or carbon/carbon.
The prior art system described above is a functional and effective thermal protection system. However, the TPS should also be minimized for weight and volume efficiency. If less weight and volume are used for the TPS, more of the missile weight and volume can be fuel or payload. Volume is generally the more critical of the two factors. Insulation materials with low thermal conductivity require less material thickness, resulting in weight and volume savings, than higher conductivity insulators.
In view of the foregoing discussion, it is apparent that there currently exists the need for an improved thermal protection system for the combustion chambers of ramjet missiles which would minimize the weight and volume spent for thermal insulation. The present invention is directed towards satisfying that need.